1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a turbine blade with near wall cooling and tip sealing.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, turbine airfoils such as rotor blades and stator vanes are cooled by passing pressurized cooling air through internal cooling air passages to produce convection cooling, impingement cooling and even film cooling of the outer airfoil surface in order to allow for higher gas flow temperatures across these airfoils. Higher turbine inlet temperatures allow for higher efficiencies of the turbine and therefore of the engine. However, the highest obtainable turbine inlet temperature is limited to the metal properties of the airfoils. Improved cooling and sealing of these airfoils will allow for higher temperatures and therefore improved performance.
U.S. Pat. No. 5,702,232 issued to Moore on Dec. 30, 1997 and entitled COOLED AIRFOILS FOR A GAS TURBINE ENGINE discloses a turbine blade with pressure and suction side walls formed around a cooling air feed chamber, where the two walls include rows of radial extending cooling air passages that extend along the entire airfoil surface and provide convection cooling to the airfoil walls. FIG. 1 shows this blade cooling system. Re-supply holes connected each radial passage to the feed chamber to re-supply cooling air and produce impingement cooling of the inner wall surface. Film cooling holes also discharge film cooling air from the radial passages and onto the outer airfoil surface. The cooling air from the radial passages also discharges out from the blade tip to produce sealing against the blade outer air seal of the engine shroud.
FIG. 2 shows one disadvantage of the blade cooling radial channels of the prior art discussed above. For the blade mid-chord section cooled with the radial flow channels, the near wall radial flow channel at the tip discharge section experiences external cross flow effect. This is represented by the arrows in FIG. 1. Because of this cross flow effect, an over-temperature will occur at the locations of the blade pressure tip regions. This external cross flow effect on the near wall radial flow channel is caused by the non-uniformity of the radial channel discharge pressure profile and the blade tip leakage flow across the radial channel exit location.